Gas turbine engine

ABSTRACT

A highly efficient gas turbine engine is provided. The fan of the gas turbine engine is driven from a turbine via a gearbox, such that the fan has a lower rotational speed than the driving turbine, thereby providing efficiency gains. The efficient fan system is mated to a core that has low cooling flow requirements and/or high temperature capability, and which may have particularly low mass for a given power.

This is a Continuation of application Ser. No. 16/411,341 filed May 14,2019, which in turn claims priority to British Application No. 1813082.3filed Aug. 10, 2018. The entire disclosures of the prior applicationsare hereby incorporated by reference herein their entirety.

The present disclosure relates to an efficient gas turbine engine.Aspects of the present disclosure relate to a gas turbine having a fandriven via a gearbox and a highly efficient engine core.

The design of a gas turbine engine must balance a number of competingfactors. In general, it is desirable to minimize fuel burn and weight.However, gas turbine engines have been used and developed for manyyears, and so the underlying designs are mature. This high level ofdesign maturity means that advances in, for example, the reduction offuel burn and/or weight have been relatively small and incremental overrecent years.

It is desirable to improve the rate of development of gas turbineengines.

According to an aspect, there is provided a gas turbine engine for anaircraft comprising:

an engine core comprising:

a first turbine, a first compressor, and a first core shaft connectingthe first turbine to the first compressor;

a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor, the secondturbine, second compressor, and second core shaft being arranged torotate at a higher rotational speed than the first core shaft; the gasturbine engine further comprising:

a fan comprising a plurality of fan blades; and

a gearbox that receives an input from the first core shaft and outputsdrive to the fan so as to drive the fan at a lower rotational speed thanthe first core shaft, wherein:

the total mass of the turbine is no greater than 17% of the total drymass of the gas turbine engine.

The total mass of the turbine may be the mass of the first turbine plusthe mass of the second turbine, for example where there are no furtherturbines in the engine.

The total mass of the turbine as a percentage of the total dry mass ofthe gas turbine engine may be in a range having a lower bound of 7%, 8%,9% or 10%, and an upper bound of 13%, 14%, 15%, 16% or 17%.

The mass of the second turbine may be no greater than 7%, 8% or 9% ofthe total dry mass of the gas turbine engine.

The mass of the second turbine as a percentage of the total dry mass ofthe gas turbine engine may be in a range having a lower bound of 3%, 4%or 5% and an upper bound of 7%, 8% or 9%.

The total dry mass of the gas turbine engine may be defined as being themass of the entire gas turbine engine prior excluding fluids (such asoil and fuel) prior to installation onto an aircraft, i.e. not includinginstallation features, such as a pylon or a nacelle.

The present inventors have found that providing a gas turbine enginewith a turbine mass in the ranges defined herein—which is lower thanconventional engines having a fan which is driven from a turbine via areduction gearbox—may provide a particularly efficient gas turbineengine.

Purely by way of example, one way of achieving such a low turbine massis through the optimal use of ceramic matrix composites in a gas turbineengine having a fan which is driven from a turbine via a reductiongearbox.

The first and/or second turbine may comprise at least one ceramic matrixcomposite component. The second turbine may comprise at least oneceramic matrix composite component, which may be in the range of from 2%to 15% of the total mass of the second turbine.

Conventionally, components in a turbine section of a gas turbine enginehave been manufactured using a metal alloy, such as a nickel alloy.However, in order to achieve greater engine efficiency, it has beenfound to be desirable to increase the temperature of the core gas flowentering into the turbine from the combustor. Typically, in operation,the temperature of the gas flowing past some of the components in theturbine is near to or above the melting point of those components. Thus,in order to ensure that such components have sufficient operating life,they require significant cooling. Such cooling is typically providedusing air from the compressor that bypasses the combustor. The coolingflow that bypasses the combustor results in reduced engine efficiency,because that flow is simply compressed in the compressor and thenexpanded through the turbine.

Furthermore, in order to minimize the amount of cooling flow that isused, and thus minimize the impact on engine efficiency, the coolingflow must be used as efficiently as possible. For example, the coolingpassages used to cool such turbine components are typically intricate,requiring extensive design and complex manufacturing techniques. Thissignificantly increases the cost of the gas turbine engine.

Still further, the cooling system itself adds mass to the engine.

Selective use of ceramic matrix composites (CMCs) in its turbine may beadvantageous. For example, CMC use may not actually be appropriate inall areas. Through this understanding, the inventors have derived theoptimum level of CMC use in the turbine to be in the claimed ranges. Forexample, whilst the thermal capability of CMCs—which is typically higherthan their metallic counterparts—may lend itself to use in some areas,the reduced thermal conductivity of CMCs (compared to an equivalentmetallic component) means that they may not be suitable in some otherareas. Purely by way of non-limitative example, the very hottest partsof the turbine may experience temperatures that exceed even thecapability of CMCs, and thus still require a degree of cooling flow. Insuch a case, it may be more appropriate to use a metal than a CMC, dueto the greater thermal conductivity of metals potentially improving theeffectiveness of the cooling flow in removing heat from the component.

Purely by way of example, where used, the CMC may be SiC-SiC (i.e.silicon carbide fibres in a silicon carbide matrix). However, it will beappreciated that any suitable CMC may be used, and indeed the turbinemay comprise more than one composition of CMC (for example havingdifferent elements). Any suitable manufacturing method may be used forthe CMC, such as a vapour deposition process or a vapour infusionprocess.

The turbine may comprise stator vanes, rotor blades, seal segments(which together may be said to form a generally annular ring radiallyoutside the rotor blades), rotor discs (on which rotor blades areprovided), one or more radially inner casing elements and one or moreradially outer casing elements. The turbine mass may be the total massof all such turbine components.

In arrangements including CMCs, the minimum mass of ceramic matrixcomposite in the second turbine may be 1%, 2%, 3%, 4%, 5%, 6%, 7%, 8%,9% or 10% of the total mass of the second turbine. The maximum mass ofceramic matrix composite in the second turbine may be 20%, 15%, 14%,13%, 12%, 11%, 10%, 9%, 8%, 7%, 6% or 5% of the total mass of the secondturbine. The mass of ceramic matrix composite in the second turbine as apercentage of the total mass of the second turbine may be in a rangehaving any of the minimum percentages listed above as a lower bound andany compatible maximum percentage listed above as an upper bound.

The second turbine may be said to be axially upstream of the firstturbine. The first turbine may comprise at least one ceramic matrixcomposite component. In arrangements including CMCs, the mass of ceramicmatrix composite in the first and second turbines may be in the range offrom 1% to 15%, optionally 2% to 12%, of the total mass of the first andsecond turbines.

In arrangements including CMCs, the minimum mass of ceramic matrixcomposite in the first and second turbines may be 1%, 2%, 3%, 4%, 5%,6%, 7%, 8%, 9% or 10% of the total mass of the first and secondturbines. In arrangements including CMCs, the maximum mass of ceramicmatrix composite in the first and second turbines may be 20%, 15%, 14%,13%, 12%, 11%, 10%, 9%, 8%, 7%, 6% or 5% of the total mass of the firstand second turbine. The mass of ceramic matrix composite in the firstand second turbines as a percentage of the total mass of the first andsecond turbines may be in a range having any of the minimum percentageslisted above as a lower bound and any compatible maximum percentagelisted above as an upper bound.

As noted above, the percentages of CMCs used in the turbine describedand claimed herein are based on insight into the most appropriatecomponents for which to use CMCs, taking into account, inter alia, thetemperature variation though the turbine. Non-limitative examples areprovided below of metallic and CMC components in the gas turbine engine

The turbine may comprise at least one row of stator vanes. The mostaxially upstream row of stator vanes may be metallic. Alternatively, themost axially upstream row of stator vanes may be CMC. The most axiallyupstream row of stator vanes may be directly downstream of thecombustor. For example, there may be no rotor blades between thecombustor and the stator vanes.

The terms “upstream” and “downstream” are used herein in theconventional manner, i.e. with respect to the flow through the engine innormal use. Thus, for example, the compressor and combustor are in theupstream direction relative to the turbine.

The turbine may comprise at least one row of rotor blades. The mostaxially upstream row of rotor blades may be metallic. Alternatively, themost axially upstream row of rotor blades may be CMC. The most axiallyupstream row of rotor blades may be directly downstream of the mostaxially upstream row of stator vanes.

The most axially upstream row of rotor blades and/or the most axiallyupstream row of stator vanes may comprise one or more internal coolingpassages and/or film cooling holes, for example where the blades and/orvanes are metallic. Such internal cooling passages and/or film coolingholes may be supplied with cooling flow from the compressor that hasbypassed the combustor.

A CMC component may or may not be provided with internal coolingpassages and/or film cooling holes.

The most axially upstream row of rotor blades in the turbine may be apart of the second turbine. The most axially upstream row of statorvanes in the turbine may be a part of the second turbine.

The most axially upstream row of rotor blades in the turbine may beradially surrounded by seal segments. Such seal segments may comprise aceramic matrix composite.

In general, the seal segments may form the radially outer boundary(which may be annular and/or frusto-conical) inside which the turbineblades rotate in use. The radially outer tips of the turbine blades maybe adjacent the radially inner surface of the seal segments.

The turbine may comprise at least two rows of stator vanes. The secondmost axially upstream row of stator vanes (which may be directly axiallydownstream of the upstream most row of rotor blades) may comprise aceramic matrix composite.

The turbine may comprise at least two rows of rotor blades. The secondmost axially upstream row of rotor blades may comprise a ceramic matrixcomposite.

The second most axially upstream row of rotor blades in the turbine maybe a part of the second turbine. The second most axially upstream row ofstator vanes in the turbine may be a part of the second turbine.

The second most axially upstream row of rotor blades may be radiallysurrounded by ceramic matrix composite seal segments.

The second turbine may comprise any number of stator vane rows (forexample 1, 2, 3, 4, 5 or 6), and one or more of which may comprise aceramic matrix composite. The second turbine may comprise any number ofrotor blade rows and/or surrounding seal segments (for example 1, 2, 3,4, 5 or 6), and one or more of which may comprise a ceramic matrixcomposite.

The axially most upstream row of stator vanes in the first turbine(which may be directly downstream of the axially most downstream row ofrotor blades in the second turbine) may comprise a ceramic matrixcomposite.

The axially most upstream row of rotor blades in the first turbine maycomprise a ceramic matrix composite. The axially most upstream row ofrotor blades in the first turbine may be surrounded by ceramic matrixcomposite seal segments.

In any aspect of the present disclosure, any one or more rotor blade,stator vane or seal segment (i.e. seal portion that forms at least apart of the radially outer flow path around a row of rotor blades) thatexperiences a maximum temperature a maximum power condition at which theengine is certified (which may be commonly known as the “red-line”condition) in the range of from 1300K to 2200K—for example in a rangehaving a lower bound of 1300K, 1400K or 1500K and an upper bound of1900K, 2000K, 2100K or 2200K—may be manufactured using a CMC. In somearrangements, most, or even all, rotor blades experiencing “red-line”temperatures within such ranges may be manufactured using a CMC. In somearrangements, most, or even all, stator vanes experiencing “red-line”temperatures within such ranges may be manufactured using a CMC. In somearrangements, most, or even all, seal segments experiencing “red-line”temperatures within such ranges may be manufactured using a CMC. Rotorblades, stator vanes and seal segments that do not experience “red-line”temperatures in such ranges may be manufactured using a metal, such as anickel alloy.

According to an aspect there is provided gas turbine engine for anaircraft comprising:

an engine core comprising:

a turbine, a combustor, and a compressor, the turbine comprising a firstturbine and a second turbine and the compressor comprising a firstcompressor and a second compressor;

a first core shaft connecting the first turbine to the first compressor;

a second core shaft connecting the second turbine to the secondcompressor, the second turbine, second compressor, and second core shaftbeing arranged to rotate at a higher rotational speed than the firstcore shaft, the gas turbine engine further comprising:

a fan comprising a plurality of fan blades; and

a gearbox that receives an input from the first core shaft and outputsdrive to the fan so as to drive the fan at a lower rotational speed thanthe first core shaft, wherein:

the second turbine comprises at least one ceramic matrix compositecomponent; and

the mass of ceramic matrix composite in the second turbine is in therange of from 2% to 15% of the total mass of the second turbine.

According to an aspect, there is provided a gas turbine engine for anaircraft comprising:

an engine core comprising:

a turbine, a compressor, and a combustor;

a fan comprising a plurality of fan blades; and

a gearbox that receives an input from the at least a part of the turbineand outputs drive to the fan so as to drive the fan at a lowerrotational speed than the first core shaft, wherein:

the turbine comprises at least one ceramic matrix composite component;and the mass of ceramic matrix composite in the turbine is in the rangeof from 1% to 15% of the total mass of the turbine, for example in therange of from 2% to 15%.

According to an aspect, there is provided a gas turbine engine for anaircraft comprising:

an engine core comprising:

a turbine, a combustor, and a compressor, the turbine comprising a firstturbine and a second turbine and the compressor comprising a firstcompressor and a second compressor;

a first core shaft connecting the first turbine to the first compressor;

a second core shaft connecting the second turbine to the secondcompressor, the second turbine, second compressor, and second core shaftbeing arranged to rotate at a higher rotational speed than the firstcore shaft, the gas turbine engine further comprising:

a bypass duct radially outside the engine core;

a fan comprising a plurality of fan blades; and

a gearbox that receives an input from the first core shaft (26) andoutputs drive to the fan so as to drive the fan at a lower rotationalspeed than the first core shaft, wherein:

part of the flow that enters the engine core bypasses the combustor andis used as turbine cooling flow to cool the turbine;

the fan diameter is greater than 225 cm and/or the turbine entrytemperature, defined as the temperature at the inlet to the most axiallyupstream turbine rotor at a maximum power condition of the gas turbineengine, is greater than 1800K; and at cruise conditions, the cooling tobypass flow efficiency ratio is less than 0.02.

The cooling to bypass efficiency ratio may be in the range of from 0.005to 0.02. The cooling to bypass efficiency ratio may be in a range havinga lower bound of 0.005, 0.006, 0.007 or 0.008, and an upper bound of0.012, 0.013, 0.014, 0.015, 0.016, 0.017, 0.018, 0.019 or 0.02.

The cooling to bypass efficiency ratio may be defined as the ratio ofthe mass flow rate of the turbine cooling flow to the mass flow rate ofthe bypass flow at engine. The ratio may be defined at engine cruiseconditions.

Such a cooling to bypass efficiency ratio—which is lower thanconventional engines—may provide a particularly efficient gas turbineengine.

According to an aspect, there is provided a gas turbine engine for anaircraft comprising:

an engine core comprising:

a first turbine, a first compressor, and a first core shaft connectingthe first turbine to the first compressor;

a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor, the secondturbine, second compressor, and second core shaft being arranged torotate at a higher rotational speed than the first core shaft, the gasturbine engine further comprising:

a fan comprising a plurality of fan blades; and

a gearbox that receives an input from the first core shaft and outputsdrive to the fan so as to drive the fan at a lower rotational speed thanthe first core shaft, wherein:

the maximum net thrust of the engine at sea level is at least 160 kN;and the normalized thrust is in the range of from 0.25 to 0.5 kN/kg.

The normalized thrust may be defined as the maximum net thrust (in kN)of the engine at sea level divided by the total mass of the turbine. Thetotal mass of the turbine may be the total mass of the first turbine andthe second turbine, for example where there are no further turbines inthe engine.

The normalized thrust may be in a range having a lower bound of 0.2,0.25 or 0.3 kN/kg and an upper bound of 0.45, 0.5 or 0.55 kN/kg.

Providing a gas turbine engine with a normalized thrust in the rangesdefined herein—which is higher than conventional engines—may provide aparticularly efficient gas turbine engine.

According to an aspect, there is provided a gas turbine engine for anaircraft comprising:

an engine core comprising:

a first turbine, a first compressor, and a first core shaft connectingthe first turbine to the first compressor;

a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor, the secondturbine, second compressor, and second core shaft being arranged torotate at a higher rotational speed than the first core shaft, the gasturbine engine further comprising:

a fan comprising a plurality of fan blades; and a gearbox that receivesan input from the first core shaft and outputs drive to the fan so as todrive the fan at a lower rotational speed than the first core shaft,wherein:

part of the flow that enters the engine core bypasses the combustor andis used as turbine cooling flow to cool the turbine;

a cooling flow requirement is defined as the ratio of the mass flow rateof the turbine cooling flow to the mass flow rate of the flow enteringthe engine core (B) at cruise conditions;

a turbine entry temperature is defined as the temperature (K) at theinlet to the most axially upstream turbine rotor in the gas turbineengine at a maximum power condition of the gas turbine engine; and

the cooling efficiency ratio, defined as the ratio between the turbineentry temperature and the cooling flow requirement, is in the range offrom 8000 to 20000 K.

The cooling efficiency ratio may be in a range having a lower bound of8000, 9000 or 10000K, and an upper bound of 18000, 20000 or 22000.

Providing a gas turbine engine with a cooling efficiency ratio in theranges defined herein—which is higher than conventional engines—mayprovide a particularly efficient gas turbine engine.

According to an aspect, there is provided a gas turbine engine for anaircraft comprising:

an engine core comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor;

a fan comprising a plurality of fan blades; and

a gearbox that receives an input from the core shaft and outputs driveto the fan so as to drive the fan at a lower rotational speed than thecore shaft, wherein:

at a maximum power condition, the ratio of the turbine entry temperature(K) to the fan speed in rpm is at least 0.7 K/rpm.

The maximum power condition may correspond to the “red-line” conditiondefined elsewhere herein.

The ratio of the turbine entry temperature (K) to the fan speed in rpmmay be in a range having a lower bound of 0.7, 0.8 or 0.9 and an upperbound of 1.5, 1.6, 1.7, 1.8, 1.9 or 2.

Providing a gas turbine engine with a ratio of the turbine entrytemperature (K) to the fan speed in rpm in the ranges definedherein—which is higher than conventional engines—may provide aparticularly efficient gas turbine engine.

According to an aspect, there is provided a gas turbine engine for anaircraft comprising:

an engine core comprising:

a first turbine, a first compressor, and a first core shaft connectingthe first turbine to the first compressor;

a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor, the secondturbine, second compressor, and second core shaft being arranged torotate at a higher rotational speed than the first core shaft, the gasturbine engine further comprising:

a fan comprising a plurality of fan blades; and

a gearbox that receives an input from the first core shaft (26) andoutputs drive to the fan so as to drive the fan at a lower rotationalspeed than the first core shaft, wherein:

a turbine entry temperature (TO_(turb_in)) is defined as the temperature(K) at the inlet to the most axially upstream turbine rotor in the gasturbine engine at a maximum power condition of the gas turbine engine;

${CS} = {{Wcomp}_{in} \cdot \frac{\sqrt{T0{comp}{\_ out}}}{P0{comp}{\_ out}}}$

a core size is defined as

where:

Wcomp_in is the mass flow rate (kg/s) at entry to the engine core;

T0comp_out is the stagnation temperature at exit to the compressor;

P0comp_out is the stagnation pressure at exit to the compressor; and

a thrust to core efficiency ratio TC is at least 1.5×10⁷ kNkg⁻¹sPa,where the thrust to

${TC} = {\left( {{Max}\mspace{14mu} {Net}\mspace{14mu} {Thrust}\mspace{14mu} {at}\mspace{14mu} {Sea}\mspace{14mu} {Level}} \right) \cdot {\frac{\sqrt{T0{turb}{\_ in}}}{CS}.}}$

core efficiency ratio is defined as

Wcomp_in may be described as being the mass flow rate at entry to thefirst compressor. T0comp_out may be described as being the stagnationtemperature at exit to the second compressor. P0comp_out may bedescribed as being the stagnation temperature at exit to the secondcompressor.

The thrust to core efficiency ratio TC may be in a range having a lowerbound of 1.5×10⁷, 1.6×10⁷, 1.7×10⁷, 1.8×10⁷, 1.9×10⁷ or 2×10⁷ kNkg⁻¹sPaand an upper bound of 3 kNkg⁻¹sPa, 3.5×10⁷ kNkg⁻¹sPa or 4 kNkg⁻¹sPa.

Providing a gas turbine engine with a thrust to core efficiency ratio inthe ranges defined herein—which is higher than conventional engines—mayprovide a particularly efficient gas turbine engine.

According to an aspect, there is provided a gas turbine engine for anaircraft comprising:

an engine core comprising:

a first turbine, a first compressor, and a first core shaft connectingthe first turbine to the first compressor;

a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor, the secondturbine, second compressor, and second core shaft being arranged torotate at a higher rotational speed than the first core shaft, the gasturbine engine further comprising:

a fan comprising a plurality of fan blades; and

a gearbox that receives an input from the first core shaft (26) andoutputs drive to the fan so as to drive the fan at a lower rotationalspeed than the first core shaft, wherein:

a turbine entry temperature (TO_(turb_in)) is defined as the temperature(K) at the inlet to the most axially upstream turbine rotor in the gasturbine engine at a maximum power condition of the gas turbine engine;

${CS} = {{Wcomp}_{in} \cdot \frac{\sqrt{T0{comp}{\_ out}}}{P0{comp}{\_ out}}}$

a core size is defined as

where:

Wcomp_in is the mass flow rate (kg/s) at entry to the engine core;

T0comp_out is the stagnation temperature at exit to the compressor;

P0comp_out is the stagnation pressure at exit to the compressor; and

a fan to core efficiency ratio FC is at least 1.9×10⁵ mkg⁻¹sPa, wherethe fan to core efficiency ratio is defined as

${FC} = {\left( {{Fan}\mspace{14mu} {Diameter}} \right) \cdot {\frac{\sqrt{T0{turb}{\_ in}}}{CS}.}}$

Wcomp_in may be described as being the mass flow rate at entry to thefirst compressor. T0comp_out may be described as being the stagnationtemperature at exit to the second compressor. P0comp_out may bedescribed as being the stagnation temperature at exit to the secondcompressor.

The fan to core efficiency ratio TC may be in a range having a lowerbound of 1.9×10⁵, 2×10⁵, or 2.1×10⁵ mkg⁻¹sPa and an upper bound of2.5×10⁵, 3×10⁵, or 3.5×10⁵ mkg⁻¹sPa.

Providing a gas turbine engine with a fan to core efficiency ratio inthe ranges defined herein—which is higher than conventional engines—mayprovide a particularly efficient gas turbine engine.

The skilled person will appreciate that except where mutually exclusive,a feature or relationship described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or relationship described herein may beapplied to any aspect and/or combined with any other feature orrelationship described herein.

By way of non-limitative example, any one or more of the followingfeatures and/or relationships disclosed herein and listed below inrelation to any aspect may be combined independently of any of the otherfeatures or relationships and/or included in any other aspect of theinvention:

Mass of ceramic matrix composite in the second turbine as a percentageof the total mass of the second turbine

Mass of ceramic matrix composite in the turbine as a whole as apercentage of the total mass of the turbine as a whole

Turbine entry temperature

Cooling to bypass flow efficiency ratio

Total mass of the turbine as a percentage of the total dry mass of thegas turbine engine

Normalized thrust of the engine

Cooling efficiency ratio

Ratio of the turbine entry temperature (K) to the fan speed in rpm

Thrust to core efficiency ratio TC

Fan to core efficiency ratio

As used herein, the turbine entry temperature, which may be referred toas TET, may be defined as the maximum temperature at entry to the mostaxially upstream rotor stage of the turbine measured at a maximum powercondition. The maximum power condition may be the maximum powercondition at which the engine is certified, and may represent themaximum temperature at that location during operation of the engine.Such a condition is commonly referred to as a “red-line” condition. Sucha condition may occur, for example, at a high thrust condition, forexample at a maximum take-off (MTO) condition. The TET (which may bereferred to as the maximum TET) in use of the engine may be particularlyhigh, for example, at least (or on the order of) any of the following:1800K, 1850K, 1900K, 1950K, 2000K, 2050K or 2100K. The maximum TET maybe in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds). Itwill be appreciated that this maximum power condition at which themaximum TET is measured is the same as the condition as that at whichthe max net thrust at sea level, or maximum thrust, (as referred toanywhere herein) is measured.

As noted elsewhere herein, the present disclosure relates to a gasturbine engine. Such a gas turbine engine may be said to comprise anengine core comprising a turbine, a combustor, a compressor, and a coreshaft connecting the turbine to the compressor. Such a gas turbineengine may comprise a fan (having fan blades) located upstream of theengine core.

As noted elsewhere herein, the gas turbine engine may comprise a gearboxthat receives an input from the core shaft and outputs drive to the fanso as to drive the fan at a lower rotational speed than the core shaft.The input to the gearbox may be directly from the core shaft, orindirectly from the core shaft, for example via a spur shaft and/orgear. The core shaft may rigidly connect the turbine and the compressor,such that the turbine and compressor rotate at the same speed (with thefan rotating at a lower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft that drives the gearbox may be afirst turbine, the compressor connected to the core shaft that drivesthe gearbox may be a first compressor, and the core shaft that drivesthe gearbox may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

The gearbox is a reduction gearbox (in that the output to the fan is alower rotational rate than the input from the core shaft). Any type ofgearbox may be used. For example, the gearbox may be a “planetary” or“star” gearbox, as described in more detail elsewhere herein. Thegearbox may have any desired reduction ratio (defined as the rotationalspeed of the input shaft divided by the rotational speed of the outputshaft), for example greater than 2.5, for example in the range of from 3to 4, for example on the order of or at least 3, 3.1, 3.2, 3.3, 3.4,3.5, 3.6, 3.7, 3.8, 3.9, 4.0, 4.1 or 4.2. The gear ratio may be, forexample, between any two of the values in the previous sentence. Purelyby way of example, the gearbox may be a “star” gearbox having a ratio inthe range of from 3.1 or 3.2 to 3.8. In some arrangements, the gearratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). These ratios may commonly be referred to as the hub-to-tipratio. The radius at the hub and the radius at the tip may both bemeasured at the leading edge (or axially forwardmost) part of the blade.The hub-to-tip ratio refers, of course, to the gas-washed portion of thefan blade, i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 225 cm, 250 cm (around 100 inches), 260 cm, 270 cm(around 105 inches), 280 cm (around 110 inches), 290 cm (around 115inches), 300 cm (around 120 inches), 310 cm, 320 cm (around 125 inches),330 cm (around 130 inches), 340 cm (around 135 inches), 350 cm, 360 cm(around 140 inches), 370 cm (around 145 inches), 380 (around 150 inches)cm, 390 cm (around 155 inches) or 400 cm. The fan diameter may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds).

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 250 cm to 300 cm (for example 250 cm to 280 cm) may be inthe range of from 1700 rpm to 2500 rpm, for example in the range of from1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100rpm. Purely by way of further non-limitative example, the rotationalspeed of the fan at cruise conditions for an engine having a fandiameter in the range of from 320 cm to 380 cm may be in the range offrom 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to1800 rpm, for example in the range of from 1400 rpm to 1600 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades on the flow results in an enthalpy rise dH of the flow. A fan tiploading may be defined as dH/U_(tip) ², where dH is the enthalpy rise(for example the 1-D average enthalpy rise) across the fan and U_(tip)is the (translational) velocity of the fan tip, for example at theleading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.28, 0.29,0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (allunits in this paragraph being Jkg⁻¹K⁻¹/(ms⁻¹)²). The fan tip loading maybe in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds).

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, or 17. The bypass ratio may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The bypass duct may besubstantially annular. The bypass duct may be radially outside the coreengine. The radially outer surface of the bypass duct may be defined bya nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds).

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹s, 105 Nkg⁻¹s, 100 Nkg⁻¹s, 95 Nkg⁻¹s, 90 Nkg⁻¹s, 85 Nkg⁻¹s or 80Nkg⁻¹s. The specific thrust may be in an inclusive range bounded by anytwo of the values in the previous sentence (i.e. the values may formupper or lower bounds). Such engines may be particularly efficient incomparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). The thrust referred to abovemay be the maximum net thrust at standard atmospheric conditions at sealevel plus 15 deg C. (ambient pressure 101.3 kPa, temperature 30 degC.), with the engine static.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a bladed disc or a bladed ring. Any suitable methodmay be used to manufacture such a bladed disc or bladed ring. Forexample, at least a part of the fan blades may be machined from a blockand/or at least part of the fan blades may be attached to the hub/discby welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26fan blades.

As used herein, cruise conditions may mean cruise conditions of anaircraft to which the gas turbine engine is attached. Such cruiseconditions may be conventionally defined as the conditions atmid-cruise, for example the conditions experienced by the aircraftand/or engine at the midpoint (in terms of time and/or distance) betweentop of climb and start of decent.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be the cruise condition. For someaircraft, the cruise conditions may be outside these ranges, for examplebelow Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions at an altitude that is in the range offrom 10000 m to 15000 m, for example in the range of from 10000 m to12000 m, for example in the range of from 10400 m to 11600 m (around38000 ft), for example in the range of from 10500 m to 11500 m, forexample in the range of from 10600 m to 11400 m, for example in therange of from 10700 m (around 35000 ft) to 11300 m, for example in therange of from 10800 m to 11200 m, for example in the range of from 10900m to 11100 m, for example on the order of 11000 m. The cruise conditionsmay correspond to standard atmospheric conditions at any given altitudein these ranges.

Purely by way of example, the cruise conditions may correspond to: aforward Mach number of 0.8; a pressure of 23000 Pa; and a temperature of−55 deg C.

As used anywhere herein, “cruise” or “cruise conditions” may mean theaerodynamic design point. Such an aerodynamic design point (or ADP) maycorrespond to the conditions (comprising, for example, one or more ofthe Mach Number, environmental conditions and thrust requirement) forwhich the fan is designed to operate. This may mean, for example, theconditions at which the fan (or gas turbine engine) is designed to haveoptimum efficiency.

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

According to an aspect, there is provided an aircraft comprising a gasturbine engine as described and/or claimed herein. The aircraftaccording to this aspect is the aircraft for which the gas turbineengine has been designed to be attached. Accordingly, the cruiseconditions according to this aspect correspond to the mid-cruise of theaircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gasturbine engine as described and/or claimed herein. The operation may beat the cruise conditions as defined elsewhere herein (for example interms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating anaircraft comprising a gas turbine engine as described and/or claimedherein. The operation according to this aspect may include (or may be)operation at the mid-cruise of the aircraft, as defined elsewhereherein.

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4 is a schematic showing an enlarged view of an upstream portion ofthe turbine of the gas turbine engine.

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14 (which may be referred to herein as a firstcompressor 14), a high-pressure compressor 15 (which may be referred toherein as a second compressor), combustion equipment 16, a high-pressureturbine 17 (which may be referred to herein as a second turbine), a lowpressure turbine 19 (which may be referred to herein as a first turbine)and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbineengine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18.The bypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low pressure turbine 19 via a shaft 26 andan epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 constrains the planet gears32 to precess around the sun gear 28 in synchronicity whilst enablingeach planet gear 32 to rotate about its own axis. The planet carrier 34is coupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Radially outwardly of the planet gears 32 andintermeshing therewith is an annulus or ring gear 38 that is coupled,via linkages 40, to a stationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2. For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20meaning that the flow through the bypass duct 22 has its own nozzle 18that is separate to and radially outside the core engine nozzle 20.However, this is not limiting, and any aspect of the present disclosuremay also apply to engines in which the flow through the bypass duct 22and the flow through the core 11 are mixed, or combined, before (orupstream of) a single nozzle, which may be referred to as a mixed flownozzle. One or both nozzles (whether mixed or split flow) may have afixed or variable area. Whilst the described example relates to aturbofan engine, the disclosure may apply, for example, to any type ofgas turbine engine, such as an open rotor (in which the fan stage is notsurrounded by a nacelle) or turboprop engine, for example.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

FIG. 4 shows a part of the turbine in greater detail. In particular,FIG. 4 shows a downstream portion of the combustor 16, the second (highpressure) turbine 17, and an upstream portion of the first (lowpressure) turbine 19. The high pressure turbine 17 is connected to thesecond core shaft 27. The low pressure turbine 19 is connected to thefirst core shaft 26.

In the illustrated example, the high pressure turbine 17 comprises, inaxial-flow series, a first (most axially upstream) stator vane row 171,a first (most axially upstream) rotor blade row 172, a second (secondmost axially upstream) stator vane row 173, and a second (second mostaxially upstream) rotor blade row 174.

The first rotor blade row 172 is connected to a rotor disc 177. Thesecond rotor blade row 174 is connected to a rotor disc 178. The tworotor discs 177, 178 are rigidly connected together by a link member179. At least one of the rotor discs (in the illustrated example thefirst rotor disc 177) is connected to the second core shaft 27 via anarm 271. Accordingly, in use, the second core shaft 27, rotor discs 177,178 and rotor blades 172, 174 all rotate together, at the samerotational speed.

The gas turbine engine 10 also comprises seal segments 175 providedradially outside the first rotor blade row 172. The gas turbine engine10 also comprises seal segments 176 provided radially outside the secondrotor blade row 174. The seal segments 175, 176 form the radially outerflow boundary (which may be referred to as the radially outer annulusline) in the region of the respective rotor blade row 172, 174, forexample over the axial extent of the tips of the rotor blades 172, 174.The seal segments 175, 176 may form a seal with the tips of the rotorblades to prevent—or at least restrict—flow passing over or past thetips of the rotor blades. The seal segments 175, 176 may be abradable bythe rotor blades. Thus, for example, the seal segments 175, 176 may beabraded by the rotor blades in use so as to form an optimal sealtherebetween. Each segment may form an annular segment or afrusto-conical segment.

In the illustrated example, the high pressure turbine 17 is a two-stagehigh pressure turbine, in that it comprises two stages of vanes andblades, each stage comprising a stator vane row followed by a rotorblade row. However, it will be appreciated that gas turbine engines 10in accordance with the present disclosure may comprise a high pressureturbine with any number of stages, for example 1, 2, 3, 4, 5 or morethan 5 stages of stator vanes and rotor blades.

The low pressure turbine 19 is provided downstream of the high pressureturbine 17. An axially most upstream row of stator vanes 191 in the lowpressure turbine 19 is provided immediately downstream of the final rowof rotor blades 174 of the high pressure turbine 17. An axially mostupstream row of rotor blades 192 in the low pressure turbine 19 isprovided immediately downstream of the axially most upstream row ofstator vanes 191. The axially most upstream row of rotor blades 192 isconnected to the first core shaft 26 via a rotor disc. In use, the rotorblades 192 of the low pressure turbine 19 drive the first core shaft 26,which in turn drives the low pressure compressor 14, and also drives—viaa gearbox 30—the fan 23.

FIG. 4 only shows an upstream portion of the low pressure turbine 19.However, it will be appreciated that downstream of the illustratedportion there may be provided further rows of stator vanes and rotorblades. For example, the low pressure turbine 19 may comprise 1, 2, 3,4, 5 or more than 5 stages of stator vanes and rotor blades. The axiallymost upstream row of rotor blades 192 are connected to one or more (notshown) downstream rotor blade rows via a linkage 199 that is connectedto the disc 197 on which the rotor blades 192 are supported.

At least a part of the high pressure turbine 17 and/or the low pressureturbine 19 comprises a CMC in the illustrated example. Purely by way ofexample, the CMC material may be silicon carbide fibres and/or a siliconcarbide matrix (SiC-SiC), although it will be appreciated that otherCMCs may be used, such as an oxide-oxide (Ox-Ox CMC material), amonolithic ceramic, and/or the like.

As noted elsewhere herein, CMCs have different properties toconventional turbine materials, such as nickel alloys. For example, CMCstypically have lower density and are able to withstand highertemperatures than metals such as nickel alloys. The present inventorshave understood that these properties can be useful in some areas of theturbine 17, 19, but other properties—such as lower thermal conductivityof CMCs compared to nickel alloys—mean that their use is not appropriatein all areas of the turbine 17, 19.

For example, depending on the type of engine (for example in terms ofarchitecture and/or maximum thrust), any one or more of the first (mostaxially upstream) stator vane row 171, first (most axially upstream)rotor blade row 172, second (second most axially upstream) stator vanerow 173, second (second most axially upstream) rotor blade row 174 andfirst or second set of seal segments 175, 176 of the high pressureturbine may be manufactured using CMCs. Components in the above listthat are not manufactured using CMCs may be manufactured using a metal,such as a nickel alloy. Optionally, in any aspect or arrangementdescribed and/or claimed herein and regardless of the number of stagesin the high pressure turbine 17, the rotor blades of each stage in thehigh pressure turbine 17 may be surrounded by seal segments, and theseal segments surrounding any one or more stage (for example all stages)may be made from a CMC.

Purely by way of non-limitative example, in the FIG. 4 arrangement, thesecond stator vane row 173, second rotor blade row 174 and first set ofseal segments 175 and second set of seal segments 176 of the highpressure turbine are manufactured using CMCs, whereas the first statorvane row 171 and the first rotor blade row 172 are manufactured using anickel alloy. In this particular example, the temperature experienced bythe first stator vane row 171 and the first rotor blade row 172 may beeven higher than that which can be tolerated by CMCs. Accordingly, forthis particular example, this means that the first stator vane row 171and the first rotor blade row 172—which experience higher temperaturesthan downstream components due to their proximity to the combustor exit16—can take advantage of the relatively high thermal conductivity of thenickel alloy so as to be cooled more effectively using cooling air(taken from the compressor, for example) which may be provided topassages running through the components.

The total mass of the high pressure turbine 17 may include the masses ofthe stator vanes 171, 173, rotor blades 172, 174, seal segments 175,176, rotor discs 177, 178, one or more radially inner casing elementsthat form the inner flow boundary 220 over the axial extent of the highpressure turbine 17, and one or more radially outer casing elements thatform the outer flow boundary 230 over the axial extent of the highpressure turbine 17.

CMCs may be used in appropriate parts of the low pressure turbine 19,although in some engines 10 their use in the low pressure turbine 19 maynot be appropriate, and thus they may not be used. Purely by way ofnon-limitative example, in the FIG. 4 arrangement, the axially mostupstream row of stator vanes 191 is manufactured using a CMC, whereasthe axially most upstream row of rotor blades 192 is manufactured usinga metal alloy (such as a nickel alloy). In this particular example, thetemperature experienced by the axially most upstream row of rotor blades192 may not be sufficiently high to benefit from the use of CMCs,although it will be appreciated that in other engines 10 in accordancewith the present disclosure, the axially most upstream row of rotorblades 192 and/or the associated seal segments 193 may be manufacturedusing CMCs. Indeed, in some engines, components (such as vanes, bladesand seals) downstream of the axially most upstream row of rotor blades192 in the low pressure turbine 19 may be manufactured using CMCs.

Any component manufactured using CMCs may also be provided with anenvironmental barrier coating (EBC). Such an EBC may cover at least thegas washed surface of such components. Such an EBC may protect the CMCfrom environmental deterioration, for example deterioration due to theeffects of water vapour. Such an EBC may be, for example ytterbiumdisilicate (Yb₂Si₂O₇), which may be applied by any suitable method, suchas air plasma spray.

As noted elsewhere herein, CMCs have a higher temperature capabilitythan conventional materials, such as metal alloys. This means thatselective use of CMCs in the turbine can mean that some components thatwould need to be cooled if they were to be made from a metal alloy donot need to be cooled because they are made from a CMC and/or somecomponents manufactured using a CMC require less cooling than if theywere to be made from a metal alloy. Additionally or alternatively,through use of CMCs it may be possible to expose some components to ahigher temperature than would otherwise be possible.

Purely by way of non-limitative example, optimizing the use of CMCs inthe engine (for example in one or more components of the turbine 17, 19as described herein) may reduce the cooling flow C requirement, whichmay result in a more efficient engine core (because less flow isbypassing the combustor), meaning that for a given amount of core power,the mass flow entering the core can be reduced and/or the size and/ormass of the turbine(s) 17, 19 can be reduced.

FIGS. 1 and 4 schematically show turbine cooling apparatus 50. Theturbine cooling apparatus extracts cooling flow C from the compressor14, 15. The cooling flow C bypasses the combustor 16. The cooling flow Cis then delivered to the high pressure turbine 17 and optionally the lowpressure turbine 19. Although the turbine cooling apparatus 50 is shownin FIGS. 1 and 4 as extracting cooling flow C from a specific positionin the high pressure compressor 15 and delivering it to a specificposition in the high pressure turbine 17, it will be appreciated thatthis is merely for ease of schematic representation, and that thecooling flow C may be extracted from any suitable locations (for examplemultiple locations in the high pressure compressor 15 and/or the lowpressure compressor 14) and delivered to any desired locations (forexample multiple locations in the high pressure turbine 17 and/or thelow pressure turbine 19).

A reduction in the amount of cooling flow C is desirable, because thecooling flow is not combusted and thus the amount of work that can beextracted from it is lower than for the flow that passes through thecombustor 16. With reference to FIG. 1, the gas turbine engine 10 has abypass ratio defined as the mass flow rate of the flow B through thebypass duct 22 divided by the mass flow rate of the flow A entering theengine core at cruise conditions. As the bypass ratio is increased—forexample to increase engine efficiency—proportionally less flow A goesthrough the core. This means that for a given size of engine and/or tobe able to withstand a given turbine entry temperature, a higherproportion of the core flow A may be required to be used as turbinecooling flow C. In this regard, as used herein, turbine entrytemperature (T0turb_in) may be the maximum stagnation temperaturemeasured at a position 60 in the gas flow path that is immediatelyupstream of the most axially upstream rotor blade row 172, i.e. at aso-called “red-line” operating condition of the engine at which theengine is certified.

A cooling to bypass efficiency ratio may be defined as the ratio of themass flow rate C of the turbine cooling flow to the mass flow rate B ofthe bypass flow at cruise conditions. Using an understanding of theconstraints and/or technologies described by way of example herein, thecooling to bypass efficiency ratio may be optimized to be as describedand/or claimed herein. Additionally or alternatively, the mass of thehigh pressure turbine 17 and/or the low pressure turbine 19 may beoptimized (for example reduced) relative to a conventional engine. Inturn, this may reduce the mass of the high pressure turbine 17 and/orthe low pressure turbine 19 as a proportion of the overall mass of thegas turbine engine 10.

Using an understanding of the constraints and/or technologies describedby way of example herein, the normalized thrust may be optimized. Inthis regard, the normalized thrust is defined as the maximum net thrustof the engine 10 at sea level divided by the total mass of the turbines17, 19 in the engine 10. The illustrated example has a high pressureturbine 17 and a low pressure turbine 19, however, it will beappreciated that where further turbines are included in the engine thetotal turbine mass includes the mass of all turbines.

As noted elsewhere herein, the optimized use of CMCs may result in areduced turbine cooling flow requirement. Additionally or alternatively,through use of CMCs it may be possible to expose some components to ahigher temperature than would otherwise be possible. This may result inthe ability to increase the turbine entry temperatures relative toengines 10 that do not include optimized use of CMCs. In this regard, ithas been found that higher turbine entry temperatures are desirable froman engine efficiency perspective.

Using an understanding of the constraints and/or technologies describedby way of example herein, the cooling efficiency ratio may be optimized.In this regard, the cooling efficiency ratio is defined as the ratiobetween the turbine entry temperature (as defined elsewhere herein) andthe cooling flow requirement. The cooling flow requirement may bedefined as the mass flow rate of the turbine cooling flow C divided bythe mass flow rate of the flow A entering the engine core at cruiseconditions.

A core size CS may be defined for the gas turbine engine 10 as:

${CS} = {{Wcomp}_{in} \cdot \frac{\sqrt{T0{comp}{\_ out}}}{P0{comp}{\_ out}}}$

where:

Wcomp_in is the mass flow rate (kg/s) at entry to the engine core, i.e.the mass flow rate of the core flow A measured at position 70 in FIG. 1;

TOcomp_out is the stagnation temperature (K) at exit to the compressor,i.e. at exit of the highest pressure compressor 15, indicated byposition 80 in FIG. 1;

P0comp_out is the stagnation pressure (Pa) at exit to the compressori.e. at exit of the highest pressure compressor 15, indicated byposition 80 in FIG. 1.

Using an understanding of the constraints and/or technologies describedby way of example herein may allow a thrust to core efficiency ratio TCand/or a fan to core efficiency ratio FC to be optimised to be in theranges described and/or claimed herein, where the thrust to coreefficiency ratio TC and the fan to core efficiency ratio FC are asdefined below (with TOturb_in being the turbine entry temperature atposition 60 shown in FIG. 4, as described above).

${TC} = {\left( {{Max}\mspace{14mu} {Net}\mspace{14mu} {Thrust}\mspace{14mu} {at}\mspace{14mu} {Sea}\mspace{14mu} {Level}} \right) \cdot \frac{\sqrt{T0{turb}{\_ in}}}{CS}}$

${FC} = {\left( {{Fan}\mspace{14mu} {Diameter}} \right) \cdot {\frac{\sqrt{T0{turb}{\_ in}}}{CS}.}}$

It will be appreciated that the understanding and/or technologydescribed and/or claimed herein results in a particularly efficient gasturbine engine 10. For example, the understanding and/or technologydescribed and/or claimed herein may provide a particularly efficient gasturbine engine 10 in which a fan 23 that is driven by a gearbox 30 iscomplemented by an optimized engine core.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features and aspects may beemployed separately or in combination with any other features and thedisclosure extends to and includes all combinations and sub-combinationsof one or more features described herein.

1. A gas turbine engine for an aircraft comprising: an engine corecomprising: a first turbine, a first compressor, and a first core shaftconnecting the first turbine to the first compressor; a second turbine,a second compressor, and a second core shaft connecting the secondturbine to the second compressor, the second turbine, second compressor,and second core shaft being arranged to rotate at a higher rotationalspeed than the first core shaft; the gas turbine engine furthercomprising: a fan comprising a plurality of fan blades; and a gearboxthat receives an input from the first core shaft and outputs drive tothe fan so as to drive the fan at a lower rotational speed than thefirst core shaft, wherein: the mass of the second turbine is in a rangeof from 4% to 8% of the total dry mass of the gas turbine engine.
 2. Agas turbine engine for an aircraft according to claim 1, wherein thetotal mass of the turbine, which includes the first turbine and thesecond turbine, is no greater than 15% of the total dry mass of the gasturbine engine.
 3. A gas turbine engine for an aircraft according toclaim 1, wherein the total mass of the turbine, which includes the firstturbine and the second turbine, is in a range of from 8% to 15% of thetotal dry mass of the gas turbine engine.
 4. A gas turbine engine for anaircraft according claim 1, wherein the mass of the second turbine is nogreater than 8% of the total dry mass of the gas turbine engine.
 5. Agas turbine engine for an aircraft according to claim 1, wherein thetotal mass of the turbine, which includes the first turbine and thesecond turbine, is in a range of from 9% to 14% of the total dry mass ofthe gas turbine engine.
 6. A gas turbine engine for an aircraftaccording to claim 1, wherein: the second turbine comprises at least oneceramic matrix composite component.
 7. A gas turbine engine for anaircraft according to claim 6, wherein the mass of ceramic matrixcomposite in the second turbine is in a range of from 2% to 15% of thetotal mass of the second turbine.
 8. A gas turbine engine for anaircraft according to claim 6, wherein: the first turbine comprises atleast one ceramic matrix composite component.
 9. A gas turbine enginefor an aircraft according to claim 1, wherein: the second turbinecomprises at least one row of stator vanes; and the most axiallyupstream row of stator vanes are metallic or ceramic matrix composite.10. A gas turbine engine for an aircraft according to claim 1, wherein:the second turbine comprises at least one row of rotor blades; and themost axially upstream row of rotor blades are metallic or ceramic matrixcomposite.
 11. A gas turbine engine for an aircraft according to claim1, wherein: the second turbine comprises at least one row of rotorblades, the most axially upstream row of rotor blades being radiallysurrounded by seal segments; and the seal segments comprise a ceramicmatrix composite.
 12. A gas turbine engine for an aircraft according toclaim 1, wherein: the second turbine comprises at least two rows ofstator vanes; and the second most axially upstream row of stator vanescomprise a ceramic matrix composite.
 13. A gas turbine engine for anaircraft according to claim 1, wherein: the second turbine comprises atleast two rows of rotor blades; and the second most axially upstream rowof rotor blades comprise a ceramic matrix composite.
 14. A gas turbineengine for an aircraft according to claim 13, wherein: the second mostaxially upstream row of rotor blades is radially surrounded by ceramicmatrix composite seal segments.
 15. A gas turbine engine for an aircraftaccording to claim 1, wherein an axially most upstream row of statorvanes in the first turbine comprise a ceramic matrix composite.
 16. Agas turbine engine for an aircraft according to claim 1, wherein anaxially most upstream row of rotor blades in the first turbine comprisea ceramic matrix composite.
 17. A gas turbine engine for an aircraftaccording to claim 1, wherein a turbine entry temperature, defined asthe temperature at the inlet to a most axially upstream turbine rotor ata maximum power condition of the gas turbine engine, is in a range offrom 1800K to 2100K.
 18. A gas turbine engine for an aircraft accordingto claim 1, wherein the fan diameter is in a range of from 225 cm to 400cm.
 19. A gas turbine engine for an aircraft according to claim 1,wherein a gear reduction ratio of the gearbox is in a range of from 3.3to
 4. 20. A gas turbine engine for an aircraft according to claim 1,wherein a maximum net thrust of the gas turbine engine at sea level isin a range of from 160 kN to 550 kN.
 21. A gas turbine engine for anaircraft according to claim 17, wherein the second turbine is cooledwith cooling air that is diverted from at least one of the first andsecond compressors.